1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as an aero engine used to power an aircraft or an industrial gas turbine engine used to produce electrical power, a turbine section includes a plurality of stages of rotor blades and stator vanes to extract the energy from the hot gas flow passing through. The engine efficiency can be improved by increasing the temperature of the hot gas flow entering the turbine. However, the inlet temperature is limited to the material properties of the first stage vanes and rotor blades. To improve the efficiency, complex internal cooling circuits have also been proposed to provide impingement and film cooling to these airfoils in order to allow for a higher gas flow temperature.
FIG. 1 shows a pressure profile for a first stage turbine blade in a prior art industrial gas turbine engine. FIG. 1 shows a graph of the pressure profile on the pressure side and on the suction side of the blade. The forward region of the pressure side surface experiences high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure. Thus, a higher pressure of cooling air is required on the pressure side than on the suction side if film cooling holes are used.
FIG. 2 shows a prior art turbine blade with a (1+5+1) serpentine flow cooling design for the first stage blade. The flow path for the 5-pass serpentine flow circuit is shown in FIG. 3. A leading edge supply channel 11 delivers cooling air to the leading edge region with a showerhead arrangement 15, a trailing edge cooling supply channel 12 delivers cooling air to exit cooling holes 16, and a supply channel or first leg 13 of the 5-pass serpentine flow circuit is positioned between the leading edge and the trailing edge channels 11 and 12 and flows forward toward the leading edge region. For a forward flowing 5-pass serpentine flow cooling circuit design used in the airfoil mid-chord region, the cooling air flows in the forward direction toward and discharges into the high hot gas side pressure section of the pressure side and the suction side through film cooling holes 17 in the 5th or last leg. In order to satisfy the back flow margin criteria (cooling air pressure for film cooling holes is higher than the external static hot gas pressure so that the hot gas does not flow into the cooling holes), a high cooling supply pressure is needed for this particular design, which induces a high leakage flow. Since the pressure of the cooling air in the 5-pass serpentine circuit decreases as it passes through the circuit toward the leading edge region, the cooling air pressure in the last or 5th leg is at its lowest. As seen from the graph in FIG. 1, the external hot gas static pressure at the last leg is higher than at any of the other legs in this serpentine circuit. Thus, the inlet pressure of the cooling air entering the first leg must be high enough so that the pressure in the last leg is high enough to prevent the hot gas flow in the last leg from entering the cooling holes.